1. Field of the Invention
The present invention relates generally to a jet aircraft engine having an improved turbine section. The invention relates more specifically to a gas turbine jet aircraft engine and a turbine section for a jet aircraft engine wherein the blades of the turbine rotor extend radially outwardly from the rotor's axis of rotation, and wherein curved manifolds are provided for directing combustion gases from the engine's combustion chamber through the turbine housing along a chordal flowpath generally perpendicular to the axis of rotation of the turbine rotor.
2. Description of Related Art
Gas-turbine jet aircraft engines were first developed in the 1920's and 1930's by Whittle. In its most basic form, the gas-turbine engine comprises an air intake, a compressor to compress the airflow from the intake, a combustion chamber where fuel is introduced to the compressed airflow and the fuel/air mixture is combusted to create a flow of hot, high-pressure combustion gasses, and a turbine to convert energy from the combustion gasses to shaft power to drive the compressor. The spent gasses exiting the turbine are expelled to the atmosphere through a nozzle or jetpipe to generate propulsive thrust. Typically, these components are arranged in axial alignment so that the working fluid passes generally linearly through the engine, from the intake to the exhaust.
A further development to gas-turbine jet aircraft engines is the turbofan engine, shown in typical form in FIG. 1. The majority of the thrust provided by a turbofan engine 10 is generated by the fan 12, which is rotationally driven by a shaft 14 within a fan cowl 16. A core cowling 18 mounted behind the fan 12 and coaxially with the fan cowl 16 forms the intake 20 for the airflow through the engine's power-generating core. This airflow first enters a compressor 22, which is typically a multistage axial compressor. Each stage of the compressor 22 typically comprises a rotating ring of rotor blades 24 within a stator ring 26. The compressor typically is driven by the drive shaft 14 coupled to the engine's turbine. Compressed air exiting the compressor 22 enters a combustion chamber 30, where fuel is introduced to form a compressed air/fuel mixture. This mixture is combusted to generate hot, high-pressure combustion gasses which are exhausted from the combustion chamber 30 to the engine's turbine section 34.
The turbine section of known gas-turbine aircraft engines commonly comprises one or more fan-like axial flow turbine rotors 36, each comprising a plurality of angularly or helically mounted turbine blades 38. Each turbine stage extracts energy from the exhaust airflow as the airflow passes axially through the blades of the turbine rotors 36. The turbine rotors, in turn, drive the shaft which powers the fan 12 and the compressor 22. Air discharged from the turbine section 34 exits the engine exhaust 40. An informative summary of existing aircraft turbine technology, and of gas-turbine aircraft engines in general, is provided by Bill Gunston, The Development of Jet and Turbine Aero Engines, pp. 39-48 (1995).
Radial flow turbines are also known, wherein high-pressure combustion gasses exiting the combustion chamber are introduced to the outer periphery of the turbine rotor and directed radially inwardly, to be discharged from a central region of the turbine. These radial flow turbines typically comprise a turbine rotor having a number of turbine blades extending in a generally spiral pattern, from adjacent the rotor's central region to the rotor's outer diameter.
Existing axial and radial flow turbines have been found, by their nature, to suffer a number of drawbacks. For example, in order to permit the high pressure combustion gasses to pass axially or radially through the turbine, a significant portion of the rotor's surface area must be open to airflow between adjacent turbine blades. Although modern multistage jet aircraft engine turbines provide greatly improved efficiencies in transferring energy from the combustion gasses to the engine's driveshaft as compared to earlier turbines, a considerable amount of energy is lost through the turbine exhaust due to the open nature of the airflow through the turbine. This lost energy results in lower horsepower-to-fuel consumption and horsepower-to-weight ratios than could be achieved were the turbine capable of more efficient energy recovery.
The open flow of combustion gasses through known axial and radial flow turbines also has been found to generate undesirable levels of noise. Moreover, because the spent combustion gasses typically exhaust through an open jetpipe downstream of the turbine, existing turbine construction has hindered efforts to provide noise and emissions control to the exhaust sections of existing jet engines. For example, the large cross-sectional area of typical jetpipes effectively prevents the utilization of known available catalytic converter or muffler equipment in the treatment of the engine's exhaust. In addition, the back-pressure which such devices typically generate would adversely affect the efficiency of known varieties of turbines.
In recent years, considerable development effort has been directed to improving the performance and efficiency of gas turbine aircraft engines. Michael Valenti, Upgrading Jet in Turbine Technology, MECHANICAL ENGINEERING, December 1995. These development efforts, however, have been primarily directed to refining existing axial or radial flow gas turbine engine geometries and improved materials of construction for conventional gas turbine engines.
Thus it can be seen that a need exists for a gas turbine jet aircraft engine comprising an improved turbine section allowing more efficient conversion of energy from the high pressure combustion gasses exiting the combustion chamber to drive the engine's driveshaft. A need also exists for a jet aircraft engine having an improved turbine enabling increased horsepower-to-fuel consumption performance. A need further exists for a jet aircraft engine turbine which generates less noise than previously existing turbines, and which permits the incorporation of noise and emissions control equipment into the engine.
It is to the provision of a jet aircraft engine and a turbine for a jet aircraft engine meeting these and other needs that the present invention is primarily directed.